1. Field of the Invention
The present invention relates to fluid reaction surfaces, and more specifically to turbine airfoils with internal cooling passages.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine burns a fuel to produce a hot gas flow which is passed through a turbine to produce mechanical work. A compressor delivers compressed air to a combustor to be mixed with a fuel to burn. The resulting hot gas flow is directed through a multiple stage turbine to convert the hot gas flow into rotation of the turbine to drive the compressor. In an Industrial Gas Turbine (IGT), the turbine shaft is also used to drive a motor such as an electrical generator. In an aero gas turbine engine, the exhaust is used to propel the airplane.
An efficiency of the gas turbine engine can be improved by operating at a higher hot gas flow temperature. However, the highest temperature in the turbine is limited to the material properties of the first stage vane and turbine blade since these are the firsts parts that the hot gas flow contacts. To allow for higher temperatures, internal cooling passages are provided within the first and second stages of the vanes and blades of the turbine. Various configurations have been proposed in order to minimize the cooling air used through the airfoils while providing for the maximum cooling effect. A multiple pass serpentine cooling flow circuit has been proposed in the past that provides improvements in both cooling air volume and cooling effectiveness.
Prior Art near wall cooling utilized in an airfoil main body is constructed with a single pass radial flow channel plus re-supply holes in conjunction with film discharge cooling holes. As a result of this cooling construction approach, spanwise and chordwise cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. In addition, a single radial channel flow is not the best method of utilizing cooling air that results in low convective cooling effectiveness.
U.S. Pat. No. 6,247,896 B1 issued to Auxier et al on Jun. 19, 2001 entitled METHOD AND APPARATUS FOR COOLING AN AIRFOIL discloses a rotor blade or stator vane for a gas turbine engine that has a plurality of micro-serpentine cooling passages formed in a wall of the airfoil and an airfoil surface formed over the passages. These passages are micro-circuits of very small diameters. Auxier et al discloses that each micro-circuit can occupy a wall surface area as great as 0.1 square inches (64.5 mm2) and with a cross sectional area of the passage segment less than 0.001 square inches (0.6 mm2). The size of these micro-circuits is so small that this invention cannot be used to cool a large blade or vane such as one used in an industrial gas turbine or IGT. The micro-circuits of the Auxier et al patent cannot provide enough cooling air flow through the passages to come near to adequately cooling the airfoil. Also, the micro-circuits of the Auxier et al patent would be very costly to make, since thousands of these tiny circuits would be needed for just one surface of the airfoil. In FIG. 3 of the Auxier et al patent, the passages appear to start from the top and flow toward the bottom due to the orientation of the Figure.
It is an object of the present invention to reduce an airfoil metal temperature of a gas turbine engine while also reducing the cooling flow requirements in order to improve turbine efficiency.
It is another object of the present invention to provide reduced airfoil metal temperature and reduced cooling flow for a large gas turbine airfoil such as one used in an industrial gas turbine engine.